Two Stage Combustor with Reformer, U.S. Patent 8,739,550
The present invention provides a combustor for an aerospace gas turbine engine comprising two stages wherein each stage defines an inlet and an exit. The second stage inlet is in fluid communication with the first stage exit such that a first flowpath is defined and it passes substantially through the second stage. A plurality of flow channel tubes is positioned within the second stage and each flow channel tube passes sealingly through a header plate positioned upstream of the second stage inlet thereby defining a second flowpath that also passes substantially through the second stage. The first flowpath exit and the second flowpath exit are positioned adjacent and proximate to one another to provide for the generation of microflames or microflame jets exiting the second stage from between and around the flow channel tube exits. The first stage of the combustor provides a gasifier and a reformer. The present invention also may comprise an igniter for further combustion of the reacted products or an external heat source for start-up. The second stage also may comprise a microflame combustor.
Etemad, Shahrokh; Baird, Benjamin; Roychoudhury, Subir; and Pfefferle, William, "Two Stage Combustor with Reformer, U.S. Patent 8,739,550" (2014). Engineering Faculty Publications. 117.
Etemad, Shahrokh, Baird, Benjamin, Roychoudhury, Subir, and Pfefferle, William "Two Stage Combustor with Reformer" U.S. Patent 8,739,550. 3 June 2014.
U.S. Patent 8,739,550 issued June 3, 2014.